EP0098363A1 - Turbine à gaz avec surveillance de la température des aubes - Google Patents

Turbine à gaz avec surveillance de la température des aubes Download PDF

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Publication number
EP0098363A1
EP0098363A1 EP83104592A EP83104592A EP0098363A1 EP 0098363 A1 EP0098363 A1 EP 0098363A1 EP 83104592 A EP83104592 A EP 83104592A EP 83104592 A EP83104592 A EP 83104592A EP 0098363 A1 EP0098363 A1 EP 0098363A1
Authority
EP
European Patent Office
Prior art keywords
turbine
cooling gas
blades
gas
cooling
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP83104592A
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German (de)
English (en)
Other versions
EP0098363B1 (fr
Inventor
Wallace B. Thomson
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Boeing North American Inc
Original Assignee
Rockwell International Corp
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Filing date
Publication date
Application filed by Rockwell International Corp filed Critical Rockwell International Corp
Publication of EP0098363A1 publication Critical patent/EP0098363A1/fr
Application granted granted Critical
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Expired legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/14—Cooling of plants of fluids in the plant, e.g. lubricant or fuel
    • F02C7/141—Cooling of plants of fluids in the plant, e.g. lubricant or fuel of working fluid
    • F02C7/143—Cooling of plants of fluids in the plant, e.g. lubricant or fuel of working fluid before or between the compressor stages
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/18—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use using the waste heat of gas-turbine plants outside the plants themselves, e.g. gas-turbine power heat plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/125—Cooling of plants by partial arc admission of the working fluid or by intermittent admission of working and cooling fluid
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention relates to a gas turbine with means for maintaining the turbine rotor blades at a low temperature. It particularly relates to a gas turbine wherein a hot gas is directed over only an annular portion of the turbine blades and a cooling gas is directed over another separate annular portion of the blades.
  • the efficiency of the gas turbine is greatly affected by the gas temperature and the efficiencies of the air compressor and the turbine.
  • Significant developments in the field of aerodynamics have greatly improved the efficiencies of the compressor and turbines such that now one of the more important keys to improving thermal efficiencies of gas turbines is by elevating the inlet gas temperatures.
  • the inlet gas temperatures are limited to those which can be withstood by the turbine materials of construction without a significant loss of structural integrity.
  • most materials utilized in commercial gas turbines are limited to a maximum temperature of about 1800 c F. (1256°K.) and generally operate at a temperature of about 1400°F. (1033°K.).
  • the present invention provides an improvement in a gas turbine which includes a housing having a hot gas inlet and a hot gas outlet, a rotatable shaft mounted in the housing and at least one turbine stage comprising a turbine rotor.
  • the turbine rotor typically includes a disc member affixed to the shaft and a plurality of turbine blades located about the periphery of the disc, each of the blades having a root section affixed to the disc and a radially outwardly extending surface terminating in a tip section.
  • the improvement comprises providing a nozzle means for directing a hot gas over only a portion or preselected number of the turbine blades. It further includes a shroud member enveloping another portion or preselected number of the turbine blades.
  • the shroud member includes an inlet and outlet for a cooling gas and a guide means is provided for directing the cooling gas radially inward or outward over the surface of the blades.
  • a deflection means directs cooling gas from the blades to the cooling gas outlet of the shroud. In operation, of course, all of the blades pass the hot gas nozzles and then pass through the shroud member where they are exposed alternately to hot gas and cooling gas, respectively.
  • the turbine further includes compressor means operably connected to the turbine shaft and having a cooling gas inlet and outlet, the outlet being in fluid communication with the shroud inlet for providing a flow of pressurized cooling gas to the shroud.
  • the pressurized cooling gas is preferably supplied at substantially the same pressure as the pressure of the hot gas which will pass over the blades of the turbine whereby leakage between the two gases is substantially minimized.
  • the guide means will provide for directing cooling gas radially downward over the tip section of the blades towards the root section and the deflection means will be located adjacent the root section of the blades. It will be appreciated, however, that the flow of cooling gas also could pass from the root section radially outward over the tip section with equal effect and minimal pressure loss. It is an essential feature of the present invention, however, that the cooling gas flow radially over the outer surfaces of the blades.
  • the gas turbine will comprise a plurality of turbine stages in axial alignment with one another and the nozzle means will provide for directing hot gas over a greater portion of the turbine blades of each successive stage located closer to the hot gas outlet of the housing.
  • a compressor means for providing cooling gas to the adjacent turbine stage.
  • the flow of pressurized cooling gas will be in a direction countercurrent to the flow of hot gas through the successive stages, as this facilitates providing cooling gas to each stage at substantially the same pressure as the hot gas flowing across that same stage.
  • the hot gas is at a relatively high pressure, for example, about 300 psia or higher, and the cooling gas must be compressed to substantially equal pressure, it is advantageous to withdraw the cooling gas from the compressor means and pass it in heat-exchange relationship with another cooling fluid to reduce the cooling gas temperature prior to introducing it into the adjacent turbine stage.
  • an axial flow gas turbine assembly 10 comprises a plurality of sections, i.e., I-IV, each section including a turbine stage and a compressor means which are operably connected to a shaft 12.
  • Gas turbine assembly 10 includes a hot gas duct 14 which extends about an increasing portion of the periphery of each successive turbine stage.
  • hot gas duct 14 extends about a peripheral arc a a and gradually expands to an even greater peripheral arc a b in section I.
  • Gas turbine assembly 10 further includes a cooling gas shroud 16 which encompasses another portion of a peripheral arc about each individual turbine stage, which arc decreases in an amount generally corresponding to the increase in the arc covered by hot gas duct 14. It will be appreciated that compression of the cooling gas will result in an increase in its temperature. In some instances the temperature increase may be such that the compressed or pressurized cooling gas leaving the compressor means has a temperature too high to maintain the turbine blades within a desired temperature range. In such instance, pressurized cooling gas from each compressor means is withdrawn through outlets 18 and cooled prior to introduction to the adjacent gas turbine stage through inlet 20.
  • the hot gas and cooling gas flow through gas turbine assembly 10 in a direction countercurrent to one another.
  • This arrangement has a particular advantage in that it minimizes the amount of ducting required. More particularly, as the hot gas passes through hot gas duct 14 and through the successive turbine stages it is expanded and cooled. Thus, the coolest and lowest pressure gas passes through the turbine stage in section I.
  • the compressor means adjacent that turbine in section I is not required to raise the pressure of the cooling gas as much as it would if it were supplying cooling gas to the turbine stage in section IV. It is, of course, preferred that the cooling gas and the hot gas passing over the same turbine stage have substantially the same static pressure to minimize any leakage between the hot gas and cooling gas portions of that turbine stage.
  • turbine stage 22 comprises a turbine rotor or disc 24 which is operably connected to shaft 12 and a plurality of turbine blades located about the periphery of rotor 24.
  • Each blade comprises a root section which is attached to rotor 24 and an outer blade surface which terminates in a radially outwardly extending tip section.
  • Located within hot gas duct 14 are a plurality of stationary turbine nozzle blades 28 for directing a flow of hot gas over a portion (preselected number) of turbine blades 26. It is a particular advantage of the present invention that the airfoil design of the turbine blade need not be compromised in the interest of cooling such as is the case where cooling air is passed internally through the turbine blades and provision must be made for such passageway.
  • Each compressor means comprises at least one and preferably a plurality of compressor wheels 30 which also are operably connected to shaft 12 and in axial alignment with turbine stage 22.
  • Each compressor wheel 30 is provided with a plurality of compressor blades 32 affixed to the outer periphery of the wheel.
  • Intermediate each compressor wheel 30 there is provided an array of compressor stators or vanes 34 for directing a compressed gas from one set of compressor blades to the next.
  • the discharge from the high pressure end of the compressor means is in fluid communication with a cooling gas manifold 36 through which the pressurized cooling gas can enter the outlet cooling duct 18 for passage to a heat exchanger (not shown) to reduce its temperature. Thereafter, the pressurized cooling gas is introduced into shroud 16 through inlet duct 20.
  • Shroud 16 includes a passageway 38 which provides fluid communication between duct 20 and a plurality of peripherally spaced, stationary cooling gas nozzle blades 40.
  • Nozzle blades 40 impart to the cooling gas a relative velocity component radially inward over the tips and outer surfaces of the rotating turbine blades 26 such that the cooling gas flows over the outer surface of the blades towards the root section of the blades.
  • a deflection means such as a beveled surface 42 for directing rapidly moving cooling gas out of shroud 16 and into the next adjacent compressor means.
  • the cooling gas be passed radially ever the outer surface of the turbine blades.
  • the reason for this is that by passing the cooling gas in such a manner the cooling effect is greatly enhanced and, further, the pressure drop is substantially reduced, thus minimizing the pumping requirements and energy losses.
  • the cooling gas is passed radially inward over the turbine blades it also is within the scope of the present invention, and sometimes may be preferred, to introduce the cooling gas adjacent the root section of the blades whereby the deflection means will direct the gas radially outward over the outer surface of the blades.
  • the gas turbine of the present invention provides many advantages which were heretofore unobtainable. More particularly, higher gas temperatures can be utilized while concurrently maintaining the temperature of the blades sufficiently low so that conventional materials of construction may be utilized. Further, since it is possible to maintain the temperature of the turbine blades well below any critical maximum value, any erosion effects resulting from particulates in the gas stream are substantially reduced. Thus, the need for having a substantially particulate-free gas stream is eliminated. Further, by providing a plurality of stages, high utilization of the energy contained in the hot gas stream is obtainable.
  • FIG. 3 therein is depicted an illustrative application of the gas turbine of the present invention which maximizes the utilization of the thermal energy contained in the hot gas introduced into the turbine.
  • a four-section gas turbine assembly 50 in which the letters "C" and "T” indicate the compressor means and turbine stages, respectively.
  • Pressurized hot gas from hot gas generator 52 is introduced into turbine assembly 50 via a conduit 54 to produce mechanical power which is utilized in any desired manner.
  • Hot gas leaving turbine assembly 50 after utilization of a substantial portion of its energy, is discharged via a duct 56 and introduced into a boiler 58 wherein a substantial amount of the remaining thermal energy is recovered to produce steam.
  • the hot gas substantially depleted of thermal energy, is discharged to the atmosphere through a stack 60.
  • the steam generated in boiler 58 is introduced into a conventional steam turbine 62 through a conduit 64 wherein the steam is expanded and cooled, converting the energy contained therein to mechanical energy which may be utilized for the generation of electrical power, driving pumps or the like.
  • the exhaust steam from steam turbine 62 is passed to a condenser 66 via a conduit 68.
  • condenser 66 the steam is cooled to a liquid state and withdrawn via a conduit 70, a pump 72 and discharged via a pipe 74, where it passes sequentially through indirect heat exchangers 76, 78, 80 and 82 and returned to boiler 58 via a conduit 84.
  • a cooling gas for example air
  • gas turbine assembly 50 A cooling gas, for example air, is introduced into gas turbine assembly 50 through an air inlet duct 86 where it passes sequentially through the four gas turbine sections.
  • the compressed cooling air is withdrawn from the compressor and introduced into its associated heat exchanger and returned to its adjacent turbine stage, as hereinbefore described.
  • the compressed gas leaving the last turbine stage in section IV passes through a duct 88 and is introduced into hot gas generator 52. Also introduced into hot gas generator 52 is a fuel through an inlet line 90.
  • the exhausted compressed cooling gas from turbine assembly 50 serves as a source of preheated combustion air for the fuel in the generation of additional hot gas, while the heat removed from the compressed pressurized cooling gas is transferred to the water introduced into boiler 58. It is seen, therefore, that in the system disclosed high thermal efficiency is obtainable.
  • any other cooling fluid used in turbine assembly 50 also can be passed in heat-exchange relationship with the feed water from boiler 58 in a similar manner. For example, when it is necessary or desired to pass a cooling fluid in indirect heat-exchange relationship with hot gas duct 14, nozzle blades 28, or both, such cooling fluid could then be passed in heat-exchange relationship with the feed water from boiler 58.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP83104592A 1982-06-14 1983-05-10 Turbine à gaz avec surveillance de la température des aubes Expired EP0098363B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US06/387,736 US4431371A (en) 1982-06-14 1982-06-14 Gas turbine with blade temperature control
US387736 1989-08-01

Publications (2)

Publication Number Publication Date
EP0098363A1 true EP0098363A1 (fr) 1984-01-18
EP0098363B1 EP0098363B1 (fr) 1988-08-03

Family

ID=23531174

Family Applications (1)

Application Number Title Priority Date Filing Date
EP83104592A Expired EP0098363B1 (fr) 1982-06-14 1983-05-10 Turbine à gaz avec surveillance de la température des aubes

Country Status (6)

Country Link
US (1) US4431371A (fr)
EP (1) EP0098363B1 (fr)
JP (1) JPS593120A (fr)
AU (1) AU559151B2 (fr)
CA (1) CA1196287A (fr)
DE (1) DE3377586D1 (fr)

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS608497A (ja) * 1983-06-29 1985-01-17 Hitachi Ltd 多段圧縮機の容量調節装置
FR2553835B1 (fr) * 1983-10-25 1986-02-28 Bertin & Cie Machine de compression d'un fluide, a plusieurs etages de compression en serie
US4784569A (en) * 1986-01-10 1988-11-15 General Electric Company Shroud means for turbine rotor blade tip clearance control
US5042970A (en) * 1989-11-28 1991-08-27 Sundstrand Corporation Fast recharge compressor
DE19643716A1 (de) * 1996-10-23 1998-04-30 Asea Brown Boveri Schaufelträger für einen Verdichter
US6398518B1 (en) * 2000-03-29 2002-06-04 Watson Cogeneration Company Method and apparatus for increasing the efficiency of a multi-stage compressor
US6752589B2 (en) * 2002-10-15 2004-06-22 General Electric Company Method and apparatus for retrofitting a steam turbine and a retrofitted steam turbine
GB2454248A (en) * 2007-11-05 2009-05-06 Siemens Ag Cooling/heating a turbomachine
EP2513431A1 (fr) * 2009-12-17 2012-10-24 Volvo Aero Corporation Agencement et procédé pour le refroidissement à circuit fermé d'un composant de moteur de turbine à gaz

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR825164A (fr) * 1937-05-24 1938-02-24 Turbine à gaz de combustion ou à gaz d'échappement, avec admission partielle
DK76136C (da) * 1950-06-29 1953-07-20 Abel Bedue Gasturbine.
CH302681A (de) * 1942-07-15 1954-10-31 Leist Karl Prof Ing Dr Verfahren zum Betrieb von Turbinen, insbesondere Gasturbinen.
US3318573A (en) * 1964-08-19 1967-05-09 Director Of Nat Aerospace Lab Apparatus for maintaining rotor disc of gas turbine engine at a low temperature
US3535873A (en) * 1967-10-24 1970-10-27 Joseph Szydlowski Gas turbine cooling devices
DE2121738A1 (de) * 1971-05-03 1972-11-09 Sontheimer, Georg, 7900 Ulm Gasturbine und Verfahren zum Betrieb derselben
US3703808A (en) * 1970-12-18 1972-11-28 Gen Electric Turbine blade tip cooling air expander
US3904307A (en) * 1974-04-10 1975-09-09 United Technologies Corp Gas generator turbine cooling scheme

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1002481A (fr) * 1946-10-09 1952-03-06 Dispositif de refroidissement pour turbines à gaz
US2951340A (en) * 1956-01-03 1960-09-06 Curtiss Wright Corp Gas turbine with control mechanism for turbine cooling air
US2925954A (en) * 1956-03-29 1960-02-23 Escher Wyss Ag Compressor group with intercooler

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR825164A (fr) * 1937-05-24 1938-02-24 Turbine à gaz de combustion ou à gaz d'échappement, avec admission partielle
CH302681A (de) * 1942-07-15 1954-10-31 Leist Karl Prof Ing Dr Verfahren zum Betrieb von Turbinen, insbesondere Gasturbinen.
DK76136C (da) * 1950-06-29 1953-07-20 Abel Bedue Gasturbine.
US3318573A (en) * 1964-08-19 1967-05-09 Director Of Nat Aerospace Lab Apparatus for maintaining rotor disc of gas turbine engine at a low temperature
US3535873A (en) * 1967-10-24 1970-10-27 Joseph Szydlowski Gas turbine cooling devices
US3703808A (en) * 1970-12-18 1972-11-28 Gen Electric Turbine blade tip cooling air expander
DE2121738A1 (de) * 1971-05-03 1972-11-09 Sontheimer, Georg, 7900 Ulm Gasturbine und Verfahren zum Betrieb derselben
US3904307A (en) * 1974-04-10 1975-09-09 United Technologies Corp Gas generator turbine cooling scheme

Also Published As

Publication number Publication date
CA1196287A (fr) 1985-11-05
DE3377586D1 (en) 1988-09-08
JPS593120A (ja) 1984-01-09
AU1521683A (en) 1983-12-22
JPH0425415B2 (fr) 1992-04-30
EP0098363B1 (fr) 1988-08-03
US4431371A (en) 1984-02-14
AU559151B2 (en) 1987-02-26

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